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' RANGER— \

^ RANGE SAFETY ANALYSIS

'AND trajectory report

Flights RA-1 and RA-2

Contract No. AF 04(647)-592

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SPACE SYSTEMS DIRECTOR #3^ fitL

LOCKHEED AIRCRAFT CORPORATION

MISSILES and SPACE DIVISION

LMSD-377728

FOREWORD

This Range Safety Analysis, including the separate Addendum
Appendix containing the trajectory information, has been prepared
by the Lockheed Missiles and Space Company under Contract
AF 04(647)-592. Authority for preparation is contained in AFSSD
Exhibit 58-1 and LMSC NASA/Agena B Program Work Statement
LMSD-446432.

This report covers only the first two Ranger flights to be launched
from Complex 12, Cape Canaveral, Florida and is submitted to
comply with the AFMTC requirement that all agencies using the
range facilities for missile flight operation must furnish data
substantiating the Range Safety capabilities of the flight articles.

Trajectory data printout listings from the IBM 7090 computer
are issued as a separate Addendum Appendix, with distribution
limited on the basis of use requirement and need-to-know.

in
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CONTENTS

FOREWORD

SECTION 1
1. 1
1. 2

1. 3

SECTION 2

2. 1
2. 2

2. 3

SECTION 3

3. 1
3. 2

3. 3

SECTION 4

4. 1
4. 2
4. 3
4.4

4. 5

SECTION 5

5. 1
5. 2
5. 3

INTRODUCTION

General

Range Safety Regulations and Requirements

Description of Vehicles

GENERAL RANGE SAFETY PROBLEMS

General

Flight Analysis Requirements

Waiver Requirements

MALFUNCTIONS

General

Atlas Failure Modes

Agena Failure Modes

FLIGHT SAFETY CRITERIA DATA

General

Tumbling Turn Rate Data

Expected Effects of a Destruct Action

Drag Estimate Plots

Miscellaneous Data

FLIGHT TERMINATION SYSTEM

General

Atlas Flight Termination System

Agena B Flight Termination System

Page
iii

1-1
1-1
1-2
1-3

2-1
2-1
2-4
2-4

3-1

3-1
3-2

3-6

4-1
4-1
4-1
4-6
4-8
4-8

5-1
5-1
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CONTENTS (Continued)

Page
APPENDIX A A-l

APPENDIX B B-l

APPENDIX C C-l

APPENDIX D D-l

REFERENCES

DISTRIBUTION

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SECTION 1
INTRODUCTION

1.1 GENERAL

The discussion and data in this report are required by the AFMTC Range
Safety Officer (RSO) in establishing safe operating limits and destruct cri-
teria for Ranger flights RA-1 and RA-2.

The purpose of these two flights is to obtain engineering test data and operat-
ing experience with the spacecraft vehicle system. The primary flight ob-
jectives are to establish the suitability of the system for subsequent hard-
landing of an instrument package on the moon and secondarily to obtain
scientific data on the space medium. A lunar trajectory is not flown on
these first two system tests.

The composite vehicle will be launched from Complex 12 at AFMTC at a pad
azimuth of 105 degrees. Shortly after liftoff, the Atlas will roll to the re-
quired trajectory of 108 degrees east of true north. The Agena first burn will
inject the vehicle and spacecraft into a nearly circular parking orbit about
the earth and Agena 2nd burn will inject the spacecraft into a highly elliptical
orbit. The trajectories are such that all impulse phases will be monitored,
from either land-based or shipboard tracking and telemetry stations.

The modified SM-65D Atlas missile will launch the composite vehicle and
carry it through the ascent trajectory to a predetermined position and alti-
tude, according to the guidance equations, where Atlas/Agena B separation
will occur. The Atlas/Agena B separation sequence begins with jettisoning
of the spacecraft shroud prior to separation, followed by separation. Agena
is then pitched down to clear the shroud. After completion of the pitchdown,

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Agena first burn ignition is initiated by the primary D-Timer. Agena first
burn cutoff is commanded by the velocity meter when circular parking orbit
conditions are attained. The Agena/ spacecraft remains in this orbit until
the point is reached where the Agena second burn is necessary to transfer
the Agena and spacecraft into the final orbit. Agena second ignition is com-
manded by the primary D-Timer which is started by the auxiliary D-Timer.
The auxiliary D-Timer is activated by a discrete command from the GE/
Burroughs Mod III-G guidance equipment during the Atlas ascent phase of
flight.

Second burn cutoff is initiated by the velocity meter after velocity required
to inject the spacecraft into an extended elliptical orbit has been achieved.
Separation of the spacecraft is followed by an Agena B retro maneuver which
will prevent it from following in the spacecraft trajectory. The unpowered
Ranger spacecraft will continue on to complete it's scientific mission.

1. 2 RANGE SAFETY REGULATIONS AND REQUIREMENTS

Since there are inherent dangers to life and property in the flight testing of
space vehicles, it is mandatory that all reasonable precautions be taken to
minimize these dangers. The Range Safety Program is designed to prevent
accidents and incidents that may result in damage to property or injury to
personnel from such operations.

Certain information and data are required by the Range in order to define
safety limits and implement range safety procedures. This includes tabu-
lations as a function of time for the predicted location of the vehicle; the
instantaneous impact points in vacuum for nominal and three sigma deviation
trajectories; and impact dispersions and maximum turn rate information.
Plots are required of the estimated drag coefficients for re-entry bodies and
re-entry pieces. Discussion is desirable on typical malfunctions and ex-
pected behavior of vehicle, expected effect of a de struct action, and other
characteristics and/or peculiarities of significance.

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The purpose of this report is to provide the needed range safety information
on behalf of the launching agency,

1. 3 DESCRIPTION OF VEHICLES

The composite flight configuration for all Ranger missions consists of three
vehicles. These are the Atlas Series D modified, the Agena B modified, and
the Ranger Spacecraft. The two booster vehicles both have several distinct
impulse phases. In the Atlas they are the booster, sustainer, and vernier
phases. In the Agena B the phases are first burn and second burn. The
composite flight vehicle is shown in Figure A-l.

1. 3. 1 Atlas

The first booster is an Atlas SM-65D modified, Serial Nos. HID and 117D,
Propellants used are liquid oxygen and RJ-1. Flight control in pitch, yaw,
and roll is accomplished by positioning the gimballed engines. Roll is
controlled by the booster and vernier engines during the first phase and by
the vernier engines during the last two impulse phases. Guidance is pro-
vided by the GE Mod III-G system. This radio-inertial guidance system
consists of a rate beacon, pulse beacon, decoder, and associated X-band
waveguide and antenna assemblies. The guidance system operates in con-
junction with related ground equipment to perform rate measuring, position
measuring, and data link functions. Position and rate information are com-
pared with pre-computed stored trajectory information by a ground computer
which generates command correction data. This data is received by the
pulse beacon where it is decoded and superimposed on the gyro-inertial and
programmer signals feeding the autopilot. The autopilot controls the engine
thrust chambers to guide the missile onto the desired flight path, thus clos-
ing the guidance loop.

Requirements for the second booster and the Lunar mission necessitates
certain structural and equipment modification in the Atlas booster. Major

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differences between a Series D R&D Atlas and a modified Atlas for the Ranger
program are as follows:

a. The forward liquid oxygen tank section skin thickness has been
increased for added rigidity, necessitated by the weight of the
Agena B and the spacecraft.

b. The standard re-entry vehicle adapter has been deleted and re-
placed with the Agena B adapter furnished by LMSD. This
adapter incorporates a self-contained Agena B separation
system.

c. The Atlas retrorockets, mountings and harness, have been
deleted from the B-l equipment pod and relocated to the Agena
adapter.

d. A section of the B-l pod-to-forward adapter wire way has been
replaced by a larger streamlined housing and mounting provision
has been made (at Station 675) for the forward rate gyros.

e. The Atlas vehicle equipment pods have been modified to reflect
electronic equipment changes and to reduce overall weight. The
upper pod is a modified B-l pod similar to that used on the
Project Mercury Atlas booster. The lower pod is extended to
accommodate GE guidance equipment relocated from the upper
pod.

f. The flight control system utilizes a more reliable transistorized
autopilot, relocated rate gyros, and a modified programmer.
The modified items in the programmer are the pitch program
and switching junction.

g. The transistorized Mod III-G radio inertial guidance set replaced
the Mod III-A because of weight and reliability considerations.

h. The Azusa system Type C transponder replaced the heavier
Type B coherent transponder.

1. 3. 2 Agena B

The orbital booster is a modified Agena B, Model 10205, Serial Nos. 6001
and 6002. It has an engine restart capability and uses IRFNA and UDMH as
propellants. Flight control is accomplished by two means after separation
from the Atlas. Pitch and yaw are controlled by positioning the gimballed
rocket engine during impulse phases and by cold gas jets during all other
phases. Roll is controlled by the cold gas system at all times. The guid-
ance system consists of an inertial reference package (IRP), a velocity

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meter, a horizon sensor, sequence timers, attitude programming circuitry,
attendant electronic assemblies, and a secondary junction box containing the
telemetry conditioning circuits. This system requires signals from the Atlas
booster to start the timer(s) and uncage the inertial reference package prior
to separation from the Atlas. Following separation, this subsystem takes
over and provides the guidance necessary to establish the desired trajectory.

1.3.3 Ranger Spacecraft

The Ranger spacecraft, Figure B-l, has been developed by the Jet Propulsion
Laboratory specifically for the Lunar missions. RA-1 and RA-2 do not con-
tain a propulsion system. After injection, maneuvers are limited to attitude
control by means of cold gas jets. The spacecraft consists of the following
components; vehicle structure, superstructure with omni-antenna, electro-
static analyzer, solar panel assemblies, 4-ft parabolic antenna, battery
packs, and six modules containing various system components and scientific
experiments. A protective shroud covers the payload during the Atlas boost
phase. Components having the highest density and structural integrity are
the six modules and battery pack. A weight breakdown of spacecraft modules
appears in Table 1-1.

Table 1-1
RANGER SPACECRAFT HIGH DENSITY COMPONENT MODULES

Principal Components Approximate Weight

Box one including 28-Ib battery 60 lbs

Box two 37

Box three 35

Box four 28

Box five 36

Box six Ul

Structure including 121-lb "battery 232

Spacecraft total » . • « » 671 lbs
Contingency .»•.*.. 15 lbs

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SECTION 2
GENERAL RANGE SAFETY PROBLEMS

2. 1 GENERAL

In general the range safety problem is restricted to that portion of the flight
that occurs between first motion and orbital velocity. Once orbital velocity
has been achieved and thrust has terminated, the instantaneous impact point
(IIP) leaves the surface of the earth and becomes indeterminate. The poten-
tial hazard to any area overflown by an orbiting vehicle then theoretically
becomes nil. Therefore, in terms of range safety it is not necessary to con-
sider those portions of flight that occur after the IIP becomes indeterminate.

Recent studies have shown that survival of hardware is possible after orbital
decay and therefore a re-entry vehicle can become a safety problem. It is
difficult, however, to predict the point of impact once an object has achieved
orbital velocity. While certain malfunctions could prevent second ignition
and cause such a problem, it is not considered to be within the range safety
domain. Re-entry of the Agena B after second burn is also in this category.

It is conceivable, but most improbable, that the flight vehicle could assume
an attitude such that the second burn impulse would cause premature re-entry.
An extended first burn would cause an elliptical orbit if the flight vehicle at-
titude has not deviated excessively from the programmed values. Re-entry
will result if the attitude has deviated excessively.

Re-entry of the flight vehicle due to a malfunction after Atlas sustainer cut-
off would be expected to cause breakup similar to that of a destruct action.
The extent of breakup and the dispersion of the re-entry pieces would be a
function of the tangential velocity at the time of malfunction. The Range
Safety Analysis for Ranger* considers two velocity conditions during the

Siegler Corporation, Inglewood, Calif., SSTG TR 61-48, NASA Ranger Program Range Safety Analysis,
10 March 1961, Secret

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Agena B first burn phase. They are: first, a premature thrust termination
just before orbital velocity is reached, and second, a premature thrust ter-
mination a few seconds after first Agena B ignition. The first condition will
generate maximum re-entry velocity and minimum re-entry angle while the
second condition gives minimum re-entry velocity and a large re-entry angle.
Large W/CjA pieces were found to impact relatively close to the vacuum
impact point, while small W/C^A pieces were found to fall farther back-
range. For the first condition, the impact point for the piece going the
farthest distance was approximately 58 nm short of the vacuum impact point
while those going the shortest distance impact approximately 165 nm short
of vacuum impact point. This gives a dispersion of approximately 107 nm in
an 11,000 foot corridor cross range. The dispersion for the second condition
was approximately 30 nm.

The mean point of impact is influenced by three factors. They are the initial
position of the vehicle at the time of the malfunction, the nature of the mal-
function, and the prevailing wind during re-entry. The previously cited
analysis considers these factors along with the lethal area of the re-entry
pieces, the reliability of the flight articles, and the population and geography
of the areas hazarded by overflight. This analysis computes the impact
probability (PJ and the kill probability (P K ) for three exit azimuths which
have a flight altitude similar (93. 7 nm vs 100 nm) to the RA-1 and RA-2
flights. These probabilities for the 108 degree azimuth are tabulated below.
They are considered to be valid for the 100 nm altitude. The islands haz-
arded by overflight on this azimuth are Rochedas Sao Pedro e Sao Paulo,
Ascension, and St. Helena. Probabilities were also computed as special
cases for Antigua on a 108 degree azimuth, and for South America on a
114 degree azimuth. Since the numbers derived for these special cases are
insignificant when compared with the numbers tabulated below, they are not
included.

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Table 2-1
SUMMARY OF Pj AND P R FOR A 108 DEGREE EXIT AZIMUTH

-5 -10

P T x 10 ? P„ x 10

Africa 5.98 1,0^3.29

Islands 0A5 0.91

Total 6.^3 l,0M+.20

During the period from shortly after launch to Atlas sustainer cutoff, the
flight vehicle will not hazard a land area on the 108 degree azimuth since
the IIP occurs on water. Moreover, trajectory deviations considerably in
excess of those expected will not endanger a land area for the same reason.

Both the first and second boosters have a de struct capability through the
Atlas Range Safety Command subsystem up to the point of Agena separation.
In the event of an abnormal flight, there are no conditions other than the
Range Safety De struct Criteria for which a premature termination of flight
is acceptable. This ensures receipt of the maximum possible data in order
to determine the cause of such abnormal performance.

During launch the pad and immediate environs are endangered until the IIP
moves over water. In the event of a malfunction this IIP can endanger land
areas uprange and crossrange as well. Destruct action during this launch
period can be as precarious as allowing impact on land. Past experience
has shown that the destruct and propellant explosion effectively destroys and
disperses most of the structural assembly. The larger pump assemblies
and electronic packages that survived impacted in a direction of the planned
trajectory, toward the ocean and away from the launch area.

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2.2 FLIGHT ANALYSIS REQUIREMENTS

The Air Force Range Safety Division AFMTC has specified the contractor
data submittal requirements for range safety analysis in a general set of
instructions*.

The significant criteria in so much as these flights are concerned is item two
of the instructions which states that trajectory data need only be presented
from launch up to a point in flight where the remaining missile stage contains
no emergency flight termination system. Therefore Appendices C and D
contains data up to Agena separation.

2. 3 WAIVER REQUIREMENTS

The Flight Range Safety Waiver requested in accordance with AFMTC
instructions* is required because:

a. The instantaneous impact point will have to cross South Africa.

b. Due to weight considerations which would compromise the
mission, a separate and independent destruct system with R-F
receivers and attendant electronic equipment is not desired.

c. Trajectory azimuths may approach or exceed normal range
safety limits.

This document, along with the impact probability study submitted previously**,
is intended to support the waiver request mentioned above. The waiver of
the Agena destruct system is based on the kill probability calculations of
reference (c)f ^therefore an attempt to analyze all re-entry burn-up conditions
is not attempted in this report.

* AFMTC/MTRSM/R M. Montgomery /65 17 (14pp), 9 June 1960

** SSTG TR 61-48, NASA Ranger Program Range Safety Analysis, Siegler Corp., lnglewood, Calif.,
10 March 1961, Secret

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SECTION 3
MALFUNCTIONS

3. 1 GENERAL

Most malfunctions are incorrect performance of the Guidance and Control or
Propulsion subsystems. Failure to separate correctly or a tracking beacon
failure are other malfunctions. A structural failure is most unlikely if the
flight vehicle is performing normally. Abnormal performance in any one
subsystem or a combination of subsystem malfunctions, can result in a
structural failure.

Breakup will certainly result if the flight vehicle does not enter orbit.
Factors contributing to vehicle breakup will vary, depending on the phase of
flight. The most obvious, and least likely, is impact with either land or
water caused by a malfunction during the early ascent phase. Control system
malfunctions which are beyond compensation by the guidance system can
cause bending failure during this period. Bending failure can also occur
during certain periods with all power off. Tension failure can result under
certain conditions when the vehicle tumbles. Aerodynamic forces will cause
breakup if the allowable angle of attack is exceeded. Aerodynamic forces
will also cause breakup during re-entry if the velocity is high enough. A
destruct action can of course cause breakup at any time that this capability
exists (launch to sustainer engine cutoff).

The malfunctions in each subsystem can be classified into major and minor
categories. From a range safety point of view a "major" malfunction is one
that will cause the flight vehicle to violate range safety restrictions while a
"minor" malfunction prevents the flight vehicle from achieving normal
trajectory.

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The discussion that follows is indicative of the wide variety of failures
which could be encountered during a flight. It is general in nature and is
not intended to be exhaustive.

3.2 ATLAS FAILURE MODES

3.2.1 Minor Failures

3.2.1.1 Incorrect Pitch Program. A pitching rate that was too low would
result in an abnormally high trajectory. In the extreme case of total ab-
sence of pitching, the missile would continue on a vertical flight path to
engine burnout or cutoff. A high pitching rate would cause the vehicle to fly
at a lower altitude than desired, and in the extreme case would continue to
pitch downward until impact while still under propulsion. Pitching 180 degrees
out of phase is another case that could occur. If the vehicle pitched
180 degrees out of phase from that which was programmed, it would start
up- range over the mainland.

Any of the forementioned cases could cause the vehicle to assume a dan-
gerous position or attitude. An incorrect pitch program could be caused by
a variety of conditions such as incorrect wiring, fluctuation of the power
supply, incorrect setting of the programmer pitch cam, and out- of- tolerance
gyro off- set.

3.2. 1.2 Incorrect Roll Program. The absence of programmed roll through
the vertical portion of the trajectory immediately after launch would result
in a trajectory azimuth of 105 degrees since the vehicle will be oriented on
the launcher for flight in that direction. Minor errors in the roll program
are corrected by the guidance system after booster cutoff.

Excessive roll of the vehicle during the vertical phase of flight would,
generally, cause the vehicle to assume a course that could endanger over-
flown areas. If this type of irregularity could not be corrected immediately
by the guidance system, the flight may have to be terminated.

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The same type of malfunctions that would cause an incorrect pitch program
could also cause an incorrect roll program. Both types of flight irregular-
ities could occur during the same period, but this is considered improbable
because of the chain of events that would have to take place simultaneously.

Closed loop tests are conducted during the countdown. This will assure that
the booster is operative as an integrated unit. In particular it will demon-
strate the proper operation of the flight programmer, guidance, autopilot,
and control systems.

3.2. 1.3 Minor Control Failures . Spurious commands to the rocket gimbal-
ling mechanism, or a malfunction of the mechanism itself, could produce
improper maneuvering. These excursions could range from a slow yaw or
pitch turn to one at the maximum turning rate. The probability of a pro-
longed turn at the maximum rate seems very low since the amount of engine
gimballing must be controlled to maintain such a turn. In addition, the
allowable angle of attack must not be exceeded at the lower altitudes.

Generally, control failures will be of the major type since the deflection
angle of the rocket engines is critical.

3.2. 1.4 Minor Propulsion Failures. Minor propulsion malfunctions are
but remotely possible. Most failures in the propulsion system are expected
to be of a major nature.

3.2.2 Major Failures

The following are typical of major failures which could be encountered
during flight.

3.2.2. 1 Major Control Failure. A major control failure could result from
spurious rocket deflection, power supply fluctuations, or anomalous dynamic
effects which are beyond the capabilities of the control system to counteract
due to the limitations on rocket engine deflections. The booster engine is

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limited to a maximum possible deflection of 4. 5 degrees, of which 2 degrees
are required to handle gusts and other disturbances. The remaining
2. 5 degrees are available to trim steady- state aerodynamic lift. As a
result, the permissible steady-state angle of attack is restricted to less
than 10 degrees during the period from 30 to 100 seconds of flight. This
trimmable angle of attack could be exceeded in less than 0. 5 second follow-
ing a major control failure.

It is estimated that the vehicle would fail structurally before it had rotated
through 90 degrees during the first 100 seconds of flight while still under
full propulsion. This could occur in less than 2 seconds in the event of
major control failure.

3.2.2.2 Major Propulsion Failure. Total loss of thrust during the first
30 seconds of flight would result in impact within or adjacent to the Cape.
Tumbling can occur at nearly any time during the first 100 seconds of flight
because the flight vehicle is aerodynamic ally unstable.

Total loss of thrust by only one of the Atlas booster thrust chambers could
send the vehicle into rapid tumbling motions, depending upon the deflection
of the functioning chamber at the time of the malfunction. The most likely
failures which could cause engine shutdown are listed below:

a. Severe shift in mixture ratio — the purpose of the propellant
utilization system is to maintain a desired mixture ratio
between the propellants as they are expended.

b. Pump or gas generator system failure — either failure will
deny propellants to the engine(s).

c. Pump cavitation — this condition will cause the pump to overspeed
which will result in a pump failure. A sufficient combination of
dynamic and static head on the pumps throughout the vehicle's
trajectory has been provided and tested on previous flights to
give a high degree of assurance that this will not occur.

The vernier engines provide roll control during main engine burn. In the
event they shut down during powered flight, the vehicle could roll, which
would result in a loss of pitch/yaw control. This loss would produce an

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erratic flight. The flight history of the Atlas shows that the possibility of
this malfunction is very improbable. Assurance of correct vernier engine
operation is obtained prior to Atlas release.

3.2.2.3 Major Structural Failure. Major structural failures could be
caused by hard-over engine deflections, or by a combination of dynamic
gust loads on the airframe. Scheduled wind soundings prior to launch will
provide wind profile data which are teletyped to Sunnyvale where they are
processed by a computer. The computer output data will be analyzed to
determine whether or not wind conditions permit a safe launch. The launch
recommendation will be transmitted to LMSD, Hanger E, at Cape Canaveral.
Additional unscheduled soundings will be taken and processed if the forecast
for the areas of interest indicate that a questionable wind condition may
exist.

A vehicle break-up study was made by Convair and three types of failures
investigated in both the pitch and yaw planes were felt to cover the extreme
possibilities. It was concluded from this study that up to approximately
120 seconds after launch, bending failure will occur under the maximum
load conditions in the yaw plane with both booster thrust chambers locked
in the extreme position § = ±4. 5 degrees. In the pitch plane from approxi-
mately 100 to 115 seconds, the vehicle could fail in bending with all power
off. Bending rather than tension failure will occur up to 120 seconds since
the angles of attack at which failure in bending occur are less than 90 degrees,
while a tension failure results from a condition in which the vehicle is
tumbling, with thrust control lost, and has passed through a 90 degree angle
of attack during its rotation.

A check was also made on the possibility of bulkhead failure due to increased
tension loads. It was concluded that a tension failure in the skin would occur
before the bulkheads failed because of the larger margin of safety for the
bulkheads.

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The Atlas test vehicles have a design structural limit of 10 g's, but it is
felt that the skin could take up to 11.5 g's in axial tension before failure.

3.2.3 Success Probability

An estimated probability of success can be made on the basis of twelve
Atlas space booster launches to date. Two were failures. The classification
of two others remain uncertain in that the cause of the flight failure could
be either the Atlas booster or the respective second booster. Two proba-
bilities can be derived: an optimistic one by assuming the two questionable
flights to be successful, and a pessimistic one by assuming the two flights
to be failures. The first probability will then be 5/6 and the second 2/3.
One de struct action was taken in these twelve space booster flights. This
is also the approximate destruct ratio for all Atlas flights.

3.3 AGENA FAILURE MODES

Due to the changing modes of vehicle operation, the effect of a subsystem
or component malfunction varies greatly with the moment of occurrence.
This discussion of Agena failure modes is arranged in the same order as
the sequence of events.

The activation of guidance and attitude control in the Agena B/ Ranger begins
at Atlas/Agena B separation; however, certain events prior to this have a
strong bearing on the behavior of the Agena B/Ranger after separation.
Malfunctions before separation in the Atlas D portion of the flight vehicle
have only indirect influence on this study and will be discussed separately.
Relative probabilities of the most likely causes of failure in the Agena B
are given in Table 3-1.

3.3.1 Failures Before Atlas/Agena B Separation

3.3. 1. 1 Primary D-Timer , The first start signal for the primary D timer
is provided by an Atlas guidance command at 287. 3 seconds followed by two

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Table 3-1
AGENA B FAILURE PROBABILITIES

Circumstances

At first Agena engine start :
Engine ignition failure
Vehicle attitude control failure
Loss of yaw control
Vehicle reference heading incorrect
Incorrect pitch attitude

During first Agena engine burn:
Premature engine cutoff
Vehicle loses pitch or yaw control
Engine locks hard over
Engine locks near zero gimbal angle

Probability

0.017723
0.00^119
0.000011
0.000002
0.000013

0.017700
0.000082
0.000088
0.000028

backup signals. Failure to start the timer will result in no operation of the
Agena B/Ranger after Atlas/Agena B separation. Due to the design of the
timer, only the following two failure modes should be recognized:

a. Synchronous motor and gear train stops.

b. Individual switch unit fails .

For item b, the consequences of the respective switch unit failures will be
discussed at the point that the unit was scheduled to operate.

3.3.1.2 Uncage Agena Gyros. The Atlas VECO signal at 295. 3 seconds also
provides the "uncage gyro" command with two backup signals following.
Failure to uncage any one gyro will result in loss of control of the corre-
sponding function (pitch, yaw, or roll) after Atlas/Agena B separation.

3.3. 1.3 Jettison Nose Fairing. Atlas guidance sends the first nose shroud
ejection signal at 296.8 seconds followed by two backup signals. Failure to
jettison the 181 pound nose shroud will not prevent the programmed attitude

3-7

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and propulsion events to bring the Agena B /Ranger into a parking orbit but
will prevent the Agena from achieving the desired final energy.

3.3.2 Failure During Atlas /Agena B Separation

The original command given by Atlas guidance is followed by two backup
signals. Since the various errors and events at separation will continue to
influence the Agena B/Ranger behavior after separation, these will be dealt
with in detail.

3.3.2. 1 Mechanical Separation. The Agena moves out of the Atlas-mounted
adapter on dual rollers engaging three guide rails spaced symmetrically
inside the adapter. Considerable clearance exists between rollers and rails.
After firing the separation pin-pullers, a 200-pound thrust is required to
initiate the separation movement, and a 3 5-pound thrust is needed to sustain
the motion. At 299. 8 seconds, Atlas guidance commands the start of the
separation cycle (followed by two backup signals). This fires the pin-puller
squibs and ignites two adapter-mounted retrorockets. Each of the latter
develops sufficient thrust to execute separation. A separation switch will
then activate the pneumatic attitude control system. Should the retrorockets
fail to ignite, the subsequent firing of the two Agena ullage rockets provides
enough thrust to push the Agena B out of the adapter.

Failure of the pin-pullers to release will not only retain the mechanical
bond between Atlas and Agena but will also prevent the attitude control
system from operating. The flight vehicle will retain the same relative
attitude as it had at Atlas VECO, subject to the angular and linear velocities
prevailing at VECO. The primary D-timer will continue to operate, as the
separation switch has no "hold-fire" circuit, and will fire the two ullage
rockets at 336.3 seconds.

It is expected that the ullage rockets, being pointed outwards, will burn
holes in the adapter wall but not cause any other damage. The Agena engine
start signal will come from the D-timer at 348. 3 seconds and a normal

3-8

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start sequence is anticipated. As the Agena engine thrust is not sufficient
to overcome the holddown bolts, it is probable that the engine jet will
destroy the front end of the Atlas and free the adapter. If an explosion
damaging the Agena does not occur, a fly-away will take place, but the roll
attitude control system will be still inoperative. The 9 degree Agena pitch-
over program will cause the engine to go hard over to correct the position;
however, the center of gravity will be behind the engine gimbal point and a
further pitch-up will occur until the Atlas falls away.

3.3.2.2 IRP Gyro Saturation. The three gyros used for pitch, yaw, and
roll control will saturate at a 12 degree error. If only one retrorocket
fires, a pitch rate of 0. 12 degree per second will be imparted to the Agena
and the time to completely separate from the Atlas will take 6 instead of
3 seconds. During separation, experience with previous flights indicate
a maximum Agena B turn rate of 1 degree per second as a result of mechan-
ical "bumping". Both these values -will cause Agena gyro errors well
below the permissible limits for the gyros used.

3.3.3 Failures After Atlas /Agena B Separation

3.3.3. 1 Coast Period . When separation is completed at about 301.8 seconds,
the Agena B/Ranger coasts for 34.5 seconds. After 11.5 seconds, a high
rate pitch-down program is initiated, lasting three seconds. A lower pitch-
down program is then begun and at the same time the horizon sensor (H/S)
is connected to the pitch and roll gyros. This latter program continues
through the first Agena engine burning period.

3.3.3.2 Time Period 301. 8 to 313. 3 Seconds . At separation, the pneu-
matic attitude control system is activated and the IRP, through the flight
control electronics and the gas jets, retains the vehicle attitude at VECO.
The following failure modes are possible:

a. Excessive or lack or yaw control

b. Excessive or lack of pitch control

c. Excessive or lack of roll control.

3-9

LOCKHEED AIRCRAFT CORPORATION _^^^^^^^^Hm MISSILES and SPACE DIVISION

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As the saturation rate of turn for the gyros in the IRP package will be
rapidly attained if any one attitude control jet opens and stays open
(0. 4 sec for roll, 2. 8 sec for pitch and yaw gyro), the vehicle will be out
of control. The same applies if complete lack of control should occur,
while a reduced rate of control will not necessarily affect the possibility of
entering the parking orbit.

3.3.3.3 Time Period 313. 3 to 316. 3 Seconds. At 313 . 3 seconds, the
primary D-timer commands a high rate (180 degree /minute) pitch-down
movement while retaining yaw and roll attitude as before. This will enable
the Agena B/Ranger to pitch down 9 degrees to the proper attitude for first
Agena engine burning. In addition to the failure modes of the previous
period, the following mode is possible.

If the primary D-timer does not command high rate pitch-down the comments
of Paragraph 3.3.3.2 are applicable.

It should also be noted that a high pitch-up attitude error will remain at the
time of the Agena engine's first start, if the primary D-timer does not com-
mand a high rate pitch-down, since the horizon sensor (H/S) will not have
sufficient gain to restore the vehicle to near horizontal attitude during the
time period 316.3 to 348.3 seconds.

3.3.3.4 Time Period 316. 3 to 336. 3 Seconds . At 31 6. 3 seconds, the
primary D-timer switches the high rate pitch-down program to a lower
rate pitch-down (3. 54 degree/minute) and connects the H/S to the pitch and
roll gyros with a control level of 3 degree/minute/degree of angular error.
The combined output of the commanded pitch- down rate and the H/S program
is calculated to keep the Agena B/Ranger at an average pitch attitude of

+ 3.2 degrees above local horizon. In addition to the failures listed in
Paragraph 3.3. 3. 2, the following modes are possible:

a. Primary D-timer does not change from high to low pitch-
down rate

b. Horizon sensor is not connected to IRP gyros.

3-10

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Failure of changeover to the lower pitch-down rate will cause the vehicle to
continue over at approximately 3 degree/ second. Should the H/S fail to give
input to the gyros, the resulting pitch- plane control would give too low a
vehicle attitude during the early period of Agena engine burning and too high
an attitude at the end of the period.

It should be recognized that a failure of certain parts in the IRP and flight
control electronics could conceivably cause temporary saturation of the yaw
gyro by creating an excessive turn rate. This would force the gyro to take
up a changed reference heading. Such an intermittent failure would then
result in the vehicle taking up a new flight azimuth. The probability of
such an event is covered in the following paragraphs.

3.3.3. 5 First Agena B Engine Burning. Agena first burn is scheduled for
the period 336. 3 to 498. 9 seconds. This time period also includes a portion
of the ullage rocket burn time which ensures engine ignition. At 33 6. 3 sec-
onds, the velocity meter is activated, and the ullage rockets ignited. At
348. 3 seconds the main Agena engine start cycle is originated by the
primary D-timer which also shifts pitch and yaw control from pneumatic to
hydraulic (engine gimballing) operation. The pitch program is unchanged
and roll control is still executed by the IRP package through the gas jets.
As the engine cutoff circuit is not activated until 487. 3 seconds, the engine
should not stop until after that time. The velocity meter cutoff point is
nominally scheduled for 498.9 seconds. The backup signal for engine cut-
off is sent by the primary D-timer at 507. 3 seconds.

The major changeover of control functions at the beginning of this period
justifies tabulating the applicable failure modes during the first Agena
engine run:

a. Engine fails to start

b. Premature engine cutoff

c. Excessive or lack of roll control (pneumatic)

d. Excessive or lack of yaw control (hydraulic)

3-11

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e. Excessive or lack of pitch control (hydraulic)

f. H/S failure

g. Primary D-timer fails to enable engine cutoff, or send backup
stop signal

h. Velocity meter provides early or late cutoff signal.

For this study, engine starting failure will result in no safety complications
because the IIP is on water. An early engine cutoff could result in unde-
sirable premature downrange impacts, while the difficulties in roll, yaw,
and pitch control would result in loss of control. A failure of the primary
D-timer to activate engine cutoff would lead to engine burning until fuel de-
pletion, which in turn would result in too high an orbital velocity and an
elliptic parking orbit. As the primary D-timer prevents the velocity meter
from sending an early engine cutoff signal before 487. 3 seconds, near park-
ing orbit velocity would be achieved. Engine cutoff at 487. 25 seconds will,
for nominal and 3a trajectories, result in Agena impacts as given in Table 3-2.
Impact occurs in a safe area in both nominal and 3a trajectories. Trajectory
parameters for a time corresponding to the "backup" engine cutoff setting,
as given in Table 3-3, are orbital for both the nominal and 3cr slow trajecto-
ries. Late engine cutoff (507. 3 seconds) would provide an orbital velocity
only slightly in excess of the desired one.

3. 3. 3. 6 Structural Failure During Ascent . Separation of the Agena B from
the Atlas during flight would be the result of a failure of the tension ties at
Station 388. 55; or, a bending failure in other areas, namely, Agena Sta-
tions 377 to 388, and in the Atlas at approximately Station 850. The most
probable cause of such a failure would be that one or more engines were
locked hard over. Separation of Station 388. 55 would automatically fire the
premature separation destruct circuit, destroying the Agena B.

3-12

LOCKHEED AIRCRAFT CORPORATION

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3-13

MISSILES and SPACE DIVISION

LMSD-377728

SECTION 4
FLIGHT SAFETY CRITERIA DATA

4. 1 GENERAL

Certain characteristics and expected behaviors of the flight vehicle are dis-
cussed in this section. Each discussion item is included to fulfill some
requirement not otherwise covered in another part of this report. They
supplement the tabulations contained in the appendices as data to establish
a safe flight criteria.

4. 2 TUMBLING TURN RATE DATA

Tumbling turn rate data for the Atlas D have been calculated' 1 ' by CV-A for
a representative flight vehicle. Calculations for tumbling turns have also
been made at LMSD for a flight vehicle very similar to the Ranger vehicle.
The results of these LMSD calculations have shown that the flight path angle
magnitudes are either equal to or somewhat less than those given in the
referenced CV-A document; and this data may therefore be used without
modification for Ranger flights RA-1 and RA-2.

It is recognized that this could result in a somewhat more restrictive de-
struct criteria than if actual rates were used. The method of calculation
of a tumbling vehicle is given below. Typical results from the LMSD calcu-
lations are compared in Figure 4-1 with th- turn rates contained in GV-A
calculations*.

* Convair- Astronautics, AZN-27-117, Angular Turn of Velocity Vector Due to Tumbling D Series Missile,
15 June 1959, Confidential

4-1

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ROLL AXIS

4.2. 1 Method of Calculation of Tumbling Vehicle

A tumbling vehicle under the action of gravity, aerodynamics, and thrust
is shown. The roll axis of the missile makes an angle 9 with the local
horizontal. The angular velocity of the roll axis is given by

6 = p + Q

(1)

where fi represents the rotation of the center of gravity about the earth's
center and Q represents the rotation of the vehicle about its center of
gravity. The first term of Equation (1) is given by the relation:

cos y

(2)

and the second term is given by the relation

Q =

6 "/t[ d

QC mQ qS + F N X + T sin SX T +: g^ X e Q

(3)

4-2

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9. 10

LMSD-377728

REFERENCE (a).

PAGE 20

GIMBAL ANGLE

= 2 DEG

GIM

BAL A

^GLE

= 4 DE

. LMSD DATA

CVAC DATA

//
//

/ /

12 3 4 5

Atg (SEC)

Figure 4-1 Turn Angle of Velocity Vector vs t for time = 24 seconds

4-3

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where

V = velocity, ft/ sec

T = thrust, lbs

S = gimbal angle, deg

a = angle of attack, deg

7 = flight path angle, deg

F^ = aerodynamic normal force, lb

g = acceleration of gravity at earth's surface, ft/ sec

I zz = pitch moment of inertia, slug-ft

X = distance between CG and CP, ft

X T = distance between CG and engine gimbal point, ft

X = distance between CG and nozzle exit plane, ft

W =• fuel flow rate, lb/ sec

CrnA = damping in pitch coefficient

S = reference area, ft

2
q = dynamic pressure, lbs/ft

radius from the center of the earth to the missile CG, ft

In order to evaluate aerodynamic coefficients in Equation (3), it is necessary
to consider high angles of attack. The coefficients will be increased over
those for the zero-lift condition due to the crossflow coefficients at high
angles of attack. The equations for center of pressure location and aero-
dynamic normal force are:

(CPJ + c N _ (CPJ

+ C

(4)

F N = C * S(C N + C N>

(5)

4-4

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where

<!).„ = Cj. cos a sin a
B

C^j = zero lift normal force coefficient

a
o

CP = center of pressure at a - deg

b B . 2 _
sin a C,

C ref co

S-r, /<-• = ratio of planform area to reference area

ref

C-. = crossflow drag coefficient
co

CP = the station at which the crossflow normal force acts
c

All items in Equations (1) through (5) are instantaneous vehicle parameters.
The solution of this set of equations gives the instantaneous turning rate in
the vertical plane. Corresponding rates for the horizontal plane are obtained
by adding the gravitational correction term:

Sj^LZ (6)

The change in 9 over any given time interval is found by integration of
Equation (1). The angle of attack, a, is given by the relation

a = tan — (7)

/J-

where

co = component of missile velocity along yaw axis
/j, = component of missile velocity along roll axis

Finally, the angle turned through by the velocity vector during the tumble is
given as:

A7 = AG - Act (8)

4-5

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4. 3 EXPECTED EFFECTS OF A DESTRUCT ACTION

4.3.1 General

It is expected that a destruct action will result in complete dispersion of the
fuels and breakup of both booster vehicles. The extent of these results will
depend on the time during flight at which the destruct action is taken. Some
of the variables involved are the amount of fuel remaining in each tank, the
velocity of the vehicle, and the density of the atmosphere. It is expected
that nearly all the fuel will be consumed through either explosion or burning.
The explosive force will depend on the quantity of fuel remaining and degree
of mixing. The quantity of fuel remaining will not change for the Agena B
booster for the duration of the command destruct capability. For the Atlas,
fuel quantity will decrease nearly linearly during each impulse phase.

The velocity of the vehicle contributes to breakup while the vehicle is in
denser atmosphere. The degree of breakup may be less if the destruct action
is taken at a later time during Atlas impulse when the remaining fuel is less
and the atmospheric density is lower. The vehicle velocity, however, will
be higher and the re-entry conditions will be more severe. This will result
in additional breakup on re-entry. In general, the extent of breakup will be
proportional to the kinetic energy of the vehicle at the time of the destruct
action.

4.3.2 Estimated Number and Size of Pieces

On the basis of these considerations it is expected that the size and quantity
of pieces that will survive to surface impact will vary with the time at which
the destruct action is taken. It is expected that a few large pieces will result
and survive from an early destruct action and that only high density pieces
having a greater structural strength are expected to survive for either a late
destruct or a normal orbital decay re-entry. The effects of a destruct action
at any other time will fall in between these extremes. The number of pieces
will increase with time because of the energy contained in the vehicle, which

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increases with time, and the reaction of the destructed vehicle with the
atmosphere during either ascent or re-entry. The number of dangerous
pieces however, will decrease with time for the same reason. An estimate
of the surviving major pieces (exclusive of skin pieces, fragments, or
molten metal pellets) and weights are tabulated below:

a. Early destruct

Surviving Piece

Atlas booster section

Atlas external equipment pod

Atlas/Agena adapter and Agena aft section

Agena battery

Ranger payload including adapter and shroud

b. Late destruct

Surviving Piece

Atlas thrust chamber gimbals
Atlas pump drive and turbine
Atlas vernier engines (each)
Agena forward equipment rack
Agena rocket motor
Ranger (several pieces)

Weight (lb)

7200
1600
1800

700

140

800

Weight (lb)

430
230
50
400
300
20-80

One destruct action has been taken on an Agena booster where the pieces
have been recovered. This was Discoverer X which was destroyed at
20, 900 feet. The second booster of this composite vehicle was an Agena A.
It is expected that a destruct action on an Agena B under the same conditions
would be very similar. The dispersion of the pieces suggests that all breakup
occurred prior to impact. The recovered pieces of the Discoverer X Agena A
are listed below:

Engine

Forward section and Payload

Hydraulic package

Nitrogen sphere

Helium sphere

Several large skin pieces

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4.3.3 Effects of Explosion on Velocity of Pieces

At all but the earliest phases of a vehicle flight, the forward velocity im-
parted to fragments by virtue of the vehicle velocity will be much greater
than that caused by the explosive force. For flight times greater than
30 seconds, the forward velocity imparted to the pieces by the explosion can
be considered to be negligible. The energy required to cause lateral disper-
sion of the pieces, normal to the vehicle velocity vector, will be supplied
solely by the forces of the explosion. The initial lateral velocity imparted to
an average size fragment by this force is estimated to be not more than
250 FPS.

4.4 DRAG ESTIMATE PLOTS

The expected effects of destruct explosion, estimates of C, versus Mach
number for pieces which,after destruct, are expected to travel maximum
and minimum distances, and the effects of explosion on velocity of pieces
for the Atlas vehicle are contained in reference (b).''~

Calculations for the Agena stage have shown that the range of C ,A/W for the
maximum and minimum distance pieces from the Agena vehicle lie within the
C ,A/W range given in reference (b) for the Atlas vehicle.

4. 5 MISCELLANEOUS DATA

The tracking equipment on board the flight vehicle which can be used for
range safety purposes is as follows:

a. Atlas telemetry transmitter

b. Atlas Azusa transponder

c. General Electric Mod III-G transponder

d. Agena "C" band transponder

* Convair Report ZC-7 136. subject "Trajectory and Aerodynamic Information for XSM-65 Missile Flights,"
26 June 1959, Secret

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The approximate time delay between activation of mainstage firing circuits
and first vehicle motion is 4. 5 seconds. First motion is defined as two inches
of movement.

The trajectory deviations expected for the most adverse wind condition for
which a launch could be expected are contained in the 3a trajectory tabula-
tions. The effects of nonstandard variations in wind, yaw gyro drift rate,
and roll gyro drift rate were included in these calculations.

The nominal duration of each impulse phase can be derived from the trajec-
tory tabulations contained in the addenda to the appendices.

The nominal duration of Atlas vernier phase is 18.43 seconds.

The maximum burn time and maximum impact range are 307. 6 seconds after
launch and 1701 nm downrange, assuming the Atlas burns to fuel exhaustion.

4-9
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LMSD-377728

SECTION 5
FLIGHT TERMINATION SYSTEM

5. 1 GENERAL

Flight termination systems are provided on each booster of the Ranger
vehicle. They provide for the termination of all flight capability from launch
to Atlas sustainer engine cutoff. Since the spacecraft contains no propellants,
a flight termination system is not considered necessary. Automatic termina-
tion is provided in the event of inadvertent separation between the first and
second boosters. These systems include a safety feature which will pre-
clude inadvertent detonation of the destruct charges during preflight check-
out and countdown on the pad. They are as follows:

a. Each system can be armed and disarmed remotely from the
blockhouse

b. Arming devices have a mechanical lockout to prevent erro-
neous arming. (A pin must be manually removed before
either destruct package can be armed)

c. Simulators are substituted for the destruct package during
prelaunch checkout.

5.2 ATLAS FLIGHT TERMINATION SYSTEM

This system is known as the Range Safety Command subsystem on the Atlas.
It is a frequency modulated system which is compatible with the ground
command destruct system at AFMTC. The system consists of two inde-
pendent command destruct paths which have in common the antenna and
coupler and the main destruct charge. Each path consists of the antenna
and coupler, command receiver, battery, arming device, and primer and
booster charges. The units of the command receiver are: broadband
converter, variable IF amplifier, fixed IF amplifier and discriminator,
audio amplifier, audio tone filters and detectors, channel relay amplifiers

5-1

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and relays, and a transistorized power supply. The arming device consists
of two relays and a switch which short-circuits the primer leads and isolates
the destructor units from receiver commands when in a safe condition. The
coupling arrangement between the antennas and the receivers is such that
the radiation pattern seen by one receiver complements the radiation
pattern seen by the other. As long as both receivers are operative, essen-
tially omnidirectional coverage is provided. If one receiver should fail
approximately 85 percent coverage is provided with an antenna gain that is
greater than 10 db below that of an isotropic radiator. Either fuel cutoff
or a de struct action will be effected when the appropriate tone combination
is received by the system. . Detailed information on this system is provided
in LMSD-447590-A, Ranger Flight Termination System , 28 February 1961,
Confidential.

5. 3 AGENA B FLIGHT TERMINATION SYSTEM

This system consists of two independent circuits which utilize the dual
command destruct signal inputs from the Atlas first booster. These circuits
and destruct charge are located in the adaptor section between the
Agena B and the Atlas. This provision will permit the destruct system to
remain with the Atlas on separation which will eliminate weight and destruct
capability during the Agena boost phase.

Dual independent premature separation destruct circuits are also provided.
Each circuit consists of independent battery, wiring, and squibs for the
initiation primer. These circuits are actuated when a lanyard releases
either of the two (2) separation switches mounted on the adaptor at the base
of the vehicle. Some components of each circuit are shared with the com-
mand destruct circuits. These circuits are disarmed and "safed" before
programmed separation of the Agena B. Detailed information on this system
is provided in LMSD-447590-A, Ranger Flight Termination System,
28 February 1961, Confidential.

5-2

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APPENDIX A

BLOCKHOUSE

OFF

ARM POWER
SWITCH

LAUNCH COORDINATORS CONSOLE
ON

TTJT
® © © ©

ARM SAFE SQUIB NO. 1 SQUIB NO. 2

ARM

safe;

SAFE

DESTRUCT SWITCH

SAFE
CHECK
SEPARATION
SWITCHES

PRESS TO TEST-

44759O-A-006

UMBILICAL JUNCTION-

HORIZON SENSORS -
VHFTM ANTENNA -

C-BAND BEACON ANTENNA -
960 MC PARASITIC ANTENNA-

DESTRUCT CHARGE -

PAD AREA

VEHICLE
ADAPTER

LOWER TM & RSC ANTENNA

LOWER GUIDANCE ANTENNA

LOWER (STUB) POD

AZUSA ANTENNA

FUEL FILL 4
DRAIN CONNECTION -

Figure A- 1 Agena-B/Atlas Vehicle

A-l

INFLIGHT
DISCONNECT
CABLE FAIRING

•RATE GYRO PACK

-FUEL TANK
•DESTRUCT UNIT

SUSTAINER
THRUST CHAMBER

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APPENDIX B

NOTE:

FOOT A, C, E HAVE TENSION/
SHEAR TIE MOUNTING POINTS

FOOT B, D, F HAVE SPRING-LOADED
SHEAR TIE MOUNTING

FOOT C IS INDEXING POINT

9A1J2 PLUG

(ELECTRICAL DISCONNECT)

SPACECRAFT
SUPPORT ASSEMBLY

9A1J1 PLUG

(ELECTRICAL DISCONNECT)

FOOTC

Figure B-l Ranger Spacecraft, Basic Structure

B-l

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APPENDIX C
VACUUM IMPACT POINTS PREMATURE ATLAS CUTOFF

C 1 ADDENDUM APPENDIX TABLES

Trajectory data printout listings from the IBM 7090 computer are issued as
a separate Addendum Appendix, with distribution limited to the list shown
under "Distribution" at the end of this document.

Tables C-l thru C-5 contain a locus of vacuum impact points for various
premature cutoff times along the nominal, 3cr maximum, 3o~ minimum, and
3cr lateral right trajectories generated using a closed-loop simulation incor-
porating the effects of the Atlas ascent guidance equations. These are the
same trajectories contained in Appendix D, Tables D- 1 thru D-4 and include
times corresponding to booster cut-off, sustainer cut-off and vernier cut-off.

Table C-l gives the expected impact points for the Atlas booster and sustainer
packages if they do not burn up upon re-entry into the atmosphere. In the
event burn up does not occur then data on the pieces will be comparable to
that given in Reference a.

Tables C-2 thru C-4 give the expected impact dispersion points about the
nominal for the Atlas booster and sustainer packages. Impact points are on
an ellipsoidal, rotating earth.

Table C-5 contains the impact points for a normal Atlas going to fuel
exhaustion which apply if normal thrust termination does not occur.

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APPENDIX D
RANGER TRAJECTORY TABLES

D. 1 ADDENDUM APPENDIX TABLES

Trajectory data printout listings from the IBM 7090 computer are issued
as a separate Addendum Appendix, with distribution limited to the list
shown under "Distribution" at the end of this document.

Tables D-l thru D-5 contain the computer listings simulating the Atlas
phase of flight for the nominal trajectory; and, for the 3o" trajectories of
maximum performance, minimum performance and lateral right deviation.
The 3t T left deviation trajectory is assumed to be symmetrical to that of
the 3a right deviation. Tables D-l thru D-3 contain the following data at
one second intervals.

a. Position coordinates (x, y s z) (ft) vs time in sec.

b. Velocity components (x, y, z ) in ft/sec vs time (sec)

V2 2 2

x + y + z (ft/sec) vs time (sec)

d. Path angle of velocity vector relative to the local horizontal
(deg) vs time (sec)

e. Altitude above earth's surface (ft) vs time (sec)

f. Ground range from launch pad to a point directly beneath the
missile (ft) vs time (sec)

g. Total weight (lbs) vs time (sec)

h. Thrust (lbs) vs time (sec)

2
i. Longitudinal acceleration (ft/sec ) vs time (sec)

The position and velocity values are referenced to an orthogonal, earth-
fixed coordinate system with origin at the launching pad. The nominal
trajectory is based upon nominal weight and engine log thrust and fuel
flow data. The 3a maximum and 3a minimum performance trajectories

D-l

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define at any time after launch the limits of normality as far as impact
range is concerned. The 3<r maximum-performance (fast, light, lofted
missile) was assumed to have high thrust, low takeoff weight, low specific
impulse, head winds, and reduced booster stage programmed pitch program.
The 3cr minimum performance (slow, heavy, low missile) has low thrust,
high take-off weight, high specific impulse, tail winds and increased
booster stage programmed pitch program.

It should be noted that Table D-5 presents the above trajectory data for a
normal Atlas going to fuel exhaustion, applies if normal thrust termination
does not occur. All of the trajectory tables in the Addendum Appendix have
been calculated using wind profile magnitudes, above which a launch would
not be attempted and therefore give the maximum expected trajectory
deviation during the first 60 to 70 seconds of flight.

D-2

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REFERENCES

(a) AZN-27-117, Angular Turn of Velocity Vector Due to Tumbling D

Series Missile , Convair- Astronautics, San Diego, Calif. , 15 June 1959

(b) ZC-7-136, Trajectory and Aerodynamic Information for XSM-65
Missile Flights , Convair-Astronautics, San Diego, Calif. ,
2 6 June 1959

(c)

SSTG TR61- 148, NASA Ranger Program Range Safety Analysis ,
Siegler Corp., Inglewood, Calif., 10 March 1961 UmmmQi)

(d)

General instructions to distribution from AFMTC /MTRSM/RM
Montgomery/ 65 17 (14 pp), Trajectory and Aerodynamic Data Require-
ments for Atlas, Titan, Atlas-Able, Atlas -Mercury, and Similar
Missiles, 9 June 196^

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DISTRIBUTION
BASIC DOCUMENT

Addressee Quantity

C ommande r 21

Air Force Space Systems Division

Air Research and Development Command

Attn: Tech. Data Center

Air Force Unit Post Office

Los Angeles 45, California

NASA Headquarters 2

1520 H Street, N. W.

Washington 25, D. C.

Attn: Lt. Cmdr. A. J. Kelley (1st copy)

NASA Headquarters

1520 H Street, N. W.

Washington 25, D. C.

Attn: W. A. Fleming (2nd copy)

George C. Marshall Space Flight Center 5

AGENA/CENTAUR Systems Office

Hunts ville, Alabama

Attn: John F. Stone, M-A & C-PR

Info: Mr. Hans Hueter, Director

Dr. K. H. Debus, Director 4

NASA Launch Operations Directorate
Titusville, Florida
Attn: Dr. A. H. Knothe

Goddard Space Flight Center 2

4555 Overlook Avenue, S. W.
Washington 25, D. C.
Attn: M. Moseson

Jet Propulsion Laboratory 2

4800 Oak Grove Drive
Pasadena, California
Attn: Clifton Cummings

Mr. H. N. Levy 1

Jet Propulsion Laboratories

COMTA

Patrick Air Force Base, Florida

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DISTRIBUTION (Continued)

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Lockheed Aircraft Corporation 1

Missiles and Space Division
Sunnyvale, California
Attn: R. E. Pace

NASA Representative

Air Force Space Systems Division 1

Air Force Unit Post Office
Los Angeles 45, California
Attn: V. G. Cummins

NASA Representative

Air Force Space Systems Division 2

Air Force Unit Post Office
Los Angeles 45, California
Attn: Maj. J. G. Albert, SSVR

Lt. Col. H. Myers 1

6555th Test Wing

AFSSD, Patrick Air Force Base, Florida

P. E. Culbertson 5

Convair Astronautics

P.O. Box 1128

San Diego 12, California

W. R. Whelan 1

Lockheed Resident Representative

Convair Astronautics

P.O. Box 1128

San Diego 12, California

W. E. Dombert 2

Burroughs Corporation
Paoli, Pennsylvania

W. Caldwell 2

AMR Test Operations Office

Burroughs Corporation

P.O. Box 105

Port Canaveral, Florida

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DISTRIBUTION
ADDENDUM APPENDIX (LMSD-377728- 1)

Addressee Quantity

C ommande r 1

Air Force Space Systems Division

Air Research and Development Command

Attn: Tech. Data Center

Air Force Unit Post Office

Los Angeles 45, California

Dr. K. H. Debus, Director 4

NASA Launch Operations Directorate
Titusville, Florida
Attn: Dr. A. H. Knothe

Air Force Space Systems Division 1

Air Force Unit Post Office
Los Angeles 45, California
Attn: Maj. J. G. Albert, SSVR

Lt. Col. H. Myers 1

6555th Test Wing

AFSSD, Patrick Air Force Base, Florida

LMSC Organizations 4

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DISTRIBUTION (Continued)

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M. M. Davis 2

Defense Systems Department
General Electric Company
Syracuse, N. Y.

J. L. Norton 2

Space Technology Laboratory

P.O. Box 95001

Los Angeles 45, California

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100

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